Efficient, low pressure ratio propulsor for gas turbine engines

ABSTRACT

A gas turbine engine includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, a first compressor coupled with the first shaft, and a second turbine coupled with the second shaft. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation of U.S. application Ser. No.15/709,567, filed Sep. 20, 2017, which is a continuation of U.S.application Ser. No. 14/695,373, filed Apr. 24, 2015, which is acontinuation-in-part of U.S. application Ser. No. 13/484,858, filed May31, 2012, which is a continuation of U.S. application Ser. No.13/176,365, filed Jul. 5, 2011.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under contract numberNAS3-01138 awarded by NASA. The government has certain rights in theinvention.

BACKGROUND

This disclosure relates to gas turbine engines and, more particularly,to an engine having a geared turbo fan architecture that is designed toefficiently operate with a high bypass ratio and a low pressure ratio.

The overall propulsive efficiency and fuel burn of a gas turbine enginedepends on many different factors, such as the design of the engine andthe resulting performance debits on the fan that propels the engine. Asan example, the fan rotates at a high rate of speed such that air passesover the blades at transonic or supersonic speeds. The fast-moving aircreates flow discontinuities or shocks that result in irreversiblepropulsive losses. Additionally, physical interaction between the fanand the air causes downstream turbulence and further losses. Althoughsome basic principles behind such losses are understood, identifying andchanging appropriate design factors to reduce such losses for a givenengine architecture has proven to be a complex and elusive task.

SUMMARY

A gas turbine engine according to an example of the present disclosureincludes a core flow passage, a bypass flow passage, and a propulsorarranged at an inlet of the bypass flow passage and the core flowpassage. The propulsor includes a row of propulsor blades. The rowincludes no more than 20 of the propulsor blades. The propulsor has apressure ratio of between about 1.2 or 1.3 and about 1.7 across thepropulsor blades.

In a further embodiment of any of the foregoing embodiments, thepressure ratio is between about 1.3 and about 1.4.

In a further embodiment of any of the foregoing embodiments, each of thepropulsor blades extends radially between a root and a tip and in achord direction between a leading edge and a trailing edge at the tip todefine a chord dimension (CD). The row of propulsor blades defines acircumferential pitch (CP) with regard to the tips. The row of propulsorblades has a solidity value (R) defined as CD/CP that is between about0.9 or 1.0 and about 1.3.

In a further embodiment of any of the foregoing embodiments, thepropulsor is coupled to be driven by a turbine through a spool, and agear assembly is coupled between the propulsor and the spool such thatrotation of the turbine drives the propulsor at a different speed thanthe spool.

In a further embodiment of any of the foregoing embodiments, thepropulsor blades include a carbon-fiber reinforced polymer matrixmaterial.

In a further embodiment of any of the foregoing embodiments, the polymerof the carbon-fiber reinforced polymer matrix material is athermoplastic polymer.

In a further embodiment of any of the foregoing embodiments, thepropulsor blades each further comprise a sheath on a leading edgethereof.

In a further embodiment of any of the foregoing embodiments, each of thepropulsor blades includes a first distinct region of carbon-fiberreinforced polymer matrix material and a second distinct region of anon-carbon-fiber reinforced polymer matrix material.

In a further embodiment of any of the foregoing embodiments, thepropulsor blades each include a distinct core that supports a skin ofcarbon-fiber reinforced polymer matrix material.

In a further embodiment of any of the foregoing embodiments, the skin ofcarbon-fiber reinforced polymer matrix material has a three-dimensionalfiber structure.

In a further embodiment of any of the foregoing embodiments, the core isformed of a metallic material.

In a further embodiment of any of the foregoing embodiments, the core isformed of a fiber-reinforced material that is different in compositionfrom the carbon-fiber reinforced polymer matrix material.

A further embodiment of any of the foregoing embodiments includes a casesurrounding the propulsor, the case including a carbon-fiber reinforcedpolymer matrix material.

A further embodiment of any of the foregoing embodiments includes a casesurrounding the propulsor. The case and the propulsor blades include acarbon-fiber reinforced polymer matrix material. The propulsor bladeseach include an airfoil body that has a distinct core that supports askin of the carbon-fiber reinforced polymer matrix material, and asheath secured on a leading edge of the airfoil body.

In a further embodiment of any of the foregoing embodiments, wherein thecarbon-fiber reinforced polymer matrix material of the propulsor bladesis different from the carbon-fiber reinforced polymer matrix material ofthe case with respect to composition.

In a further embodiment of any of the foregoing embodiments, the rowincludes no more than 17 of the propulsor blades.

In a further embodiment of any of the foregoing embodiments, thepropulsor blades each include a distinct core that supports a skin ofthe carbon-fiber reinforced polymer matrix material.

In a further embodiment of any of the foregoing embodiments, the fiberreinforced polymer matrix material of the case includes carbon fibers.

A gas turbine engine according to an example of the present disclosureincludes a core flow passage, a bypass flow passage, and a propulsorarranged at an inlet of the bypass flow passage and the core flowpassage. The propulsor includes a row of propulsor blades. The rowincludes no more than 20 of the propulsor blades and the propulsorblades include a carbon-fiber reinforced polymer matrix material. Eachof the propulsor blades extends radially between a root and a tip and ina chord direction between a leading edge and a trailing edge at the tipto define a chord dimension (CD). The row of propulsor blades defines acircumferential pitch (CP) with regard to the tips. The row of propulsorblades has a solidity value (R) defined as CD/CP that is less than about1.2 or less than about 1.1.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the disclosed examples willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 is a schematic cross-section of an embodiment of a gas turbineengine.

FIG. 2 is a perspective view of a fan section of the engine of FIG. 1.

FIG. 3 illustrates an embodiment of a carbon-fiber reinforced polymermatrix material.

FIG. 4 illustrates an embodiment of a two-dimensional woven fiberstructure.

FIG. 5 illustrates an embodiment of a three-dimensional fiber structure.

FIG. 6 is a cross-section of an embodiment of a propulsor blade that hasa distinct core and a skin of carbon-fiber reinforced polymer matrixmaterial.

FIG. 7 illustrates an embodiment of a propulsor blade that has a sheath.

FIG. 8 illustrates a portion of an embodiment of a case and propulsorblade.

FIG. 9 illustrates another embodiment of a case.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 may be a two-spool turbofan that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engine architecturesmay include a single-spool design, a three-spool design, or an openrotor design, among other systems or features.

The fan section 22 drives air along a bypass flow passage B while thecompressor section 24 drives air along a core flow passage C forcompression and communication into the combustor section 26. Althoughdepicted as a turbofan gas turbine engine, it is to be understood thatthe concepts described herein are not limited to use with turbofans andthe teachings may be applied to other types of gas turbine engines.

The engine 20 includes a low speed spool 30 and high speed spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearing systems38. The low speed spool 30 generally includes an inner shaft 40 that iscoupled with a propulsor 42, a low pressure compressor 44 and a lowpressure turbine 46. The propulsor 42 is in the fan section 22 and acase 43 surrounds the propulsor 42. The low pressure turbine 46 drivesthe propulsor 42 through the inner shaft 40 and a gear assembly 48,which allows the low speed spool 30 to drive the propulsor 42 at adifferent (e.g. lower) angular speed.

The high speed spool 32 includes an outer shaft 50 that is coupled witha high pressure compressor 52 and a high pressure turbine 54. Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis A,which is collinear with their longitudinal axes.

A core airflow in core flow passage C is compressed by the low pressurecompressor 44 then the high pressure compressor 52, mixed with the fuelin the combustor 56, and then expanded over the high pressure turbine 54and low pressure turbine 46. The turbines 54, 46 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

As shown, the propulsor 42 is arranged at an inlet 60 of the bypass flowpassage B and core flow passage C. Air flow through the bypass flowpassage B exits the engine 20 through an outlet 62 or nozzle. For agiven design of the propulsor 42, the inlet 60 and the outlet 62 of theengine 20 define a design (fan) pressure ratio with regard to an inletpressure at the inlet 60 and an outlet pressure at the outlet 62 of thebypass flow passage B. As an example, the design pressure ratio may bedetermined based upon the stagnation inlet pressure and the stagnationoutlet pressure at a design rotational speed of the engine 20. In thatregard, the engine 20 may optionally include a variable area nozzle 64within the bypass flow passage B. The variable area nozzle 64 isoperative to change a cross-sectional area 66 of the outlet 62 tothereby control the pressure ratio via changing pressure within thebypass flow passage B. The design pressure ratio may be defined with thevariable area nozzle 64 fully open or fully closed.

Referring to FIG. 2, the propulsor 42, which in this example is a fan,includes a rotor 70 having a row 72 of propulsor blades 74 that extend acircumferentially around a hub 76. Each of the propulsor blades 74extends radially outwardly from the hub 76 between a root 78 and a tip80 and in a chord direction (axially and circumferentially) between aleading edge 82 and a trailing edge 84. A chord dimension (CD) is alength between the leading edge 82 and the trailing edge 84 at the tipof each propulsor blade 74. The row 72 of propulsor blades 74 alsodefines a circumferential pitch (CP) that is equivalent to the arcdistance between the tips 80 of neighboring propulsor blades 74.

As will be described, the example propulsor 42 includes a number (N) ofthe propulsor blades 74 and a geometry that, in combination with thearchitecture of the engine 20, provides enhanced overall propulsiveefficiency by reducing performance debits of the propulsor 42.

In the illustrated example, the number N of propulsor blades in the row72 is no more than 20. In one example, the propulsor 42 includes 18 ofthe propulsor blades 74 uniformly circumferentially arranged about thehub 76. In other embodiments, the number N may be any number of bladesfrom 12-20.

The propulsor blades 74 define a solidity value with regard to the chorddimension CD and the circumferential pitch CP. The solidity value isdefined as a ratio (R) of CD/CP (i.e., CD divided by CP). Inembodiments, the solidity value of the propulsor 42 is between 0.9 or1.0 and 1.3. In further embodiments, the solidity value is from 1.1 to1.2. In additional embodiments, the solidity value is less than 1.1, andin a further example is also greater than 0.85.

Additionally, in combination with the given example solidity values, thefan 22 of the engine 20 may be designed with a particular designpressure ratio. In embodiments, the design pressure ratio may be between1.2 or 1.3 and 1.55. In a further embodiment, the design pressure ratiomay be between 1.3 and 1.4. In further examples, the design pressureratio is between 1.3 and 1.7.

The engine 20 may also be designed with a particular bypass ratio withregard to the amount of air that passes through the bypass flow passageB and the amount of air that passes through the core flow passage C. Asan example, the design bypass ratio of the engine 20 may nominally be12, or alternatively in a range of approximately 8.5 to 13.5 or 18.

The propulsor 42 also defines a ratio of N/R. In embodiments, the ratioN/R is from 9 to 20. In further embodiments, the ratio N/R is from 14 to16. The table below shows additional examples of solidity and the ratioN/R for different numbers of propulsor blades 74.

TABLE Number of Blades, Solidity and Ratio N/R Number of Blades (N)Solidity Ratio N/R 20 1.3 15.4 18 1.3 13.8 16 1.3 12.3 14 1.3 10.8 121.3 9.2 20 1.2 16.7 18 1.2 15.0 16 1.2 13.3 14 1.2 11.7 12 1.2 10.0 201.1 18.2 18 1.1 16.4 16 1.1 14.5 14 1.1 12.7 12 1.1 10.9 20 1.0 20.0 181.0 18.0 16 1.0 16.0 14 1.0 14.0 12 1.0 12.0

The disclosed ratios of N/R enhance the overall propulsive efficiencyand fuel burn of the disclosed engine 20. For instance, the disclosedratios of N/R are designed for the geared turbo fan architecture of theengine 20 that utilizes the gear assembly 48. That is, the gear assembly48 allows the propulsor 42 to rotate at a different, lower speed thanthe low speed spool 30. In combination with the variable area nozzle 64,the propulsor 42 can be designed with a large diameter and rotate at arelatively slow speed with regard to the low speed spool 30. Arelatively low speed, relatively large diameter, and the geometry thatpermits the disclosed ratios of N/R contribute to the reduction ofperformance debits, such as by lowering the speed of the air or fluidthat passes over the propulsor blades 74.

The propulsor blades 74 can include a carbon-fiber reinforced polymermatrix material, an example portion of which is depicted in FIG. 3 at86. In this example, the material 86 includes carbon fibers 86 a thatare disposed in a polymer matrix 86 b. The propulsor blades 74 can beformed exclusively of the material 86 or partially of the material 86 incombinations with alloys or other fiber-reinforced materials.

The material 86 can include a plurality of carbon fiber layers 88 thatare stacked and consolidated to form the material 86. For example, thefiber layers 88 can each have uni-directionally oriented fibers and thelayers 88 can be cross-plied. In further examples, one or more of thelayers 88 has a different fiber structure, such as but not limited to,random fiber orientation, woven, or three-dimensional. An exampletwo-dimensional woven fiber structure is depicted in FIG. 4. An examplethree-dimensional fiber structure is depicted in FIG. 5. In thisexample, the fibers 86 a are woven into sheets 90, and transverse fibers86 c bundle the sheets 90 to one another. As can be appreciated, othertwo- or three-dimensional fiber structures could alternatively oradditionally be used.

The polymer matrix 86 b can include thermoplastic polymer, thermosetpolymer, or combinations thereof. Thermoset polymers can include, butare not limited to, epoxy and phenolic. Thermoplastic polymers caninclude, but are not limited to, polyethers and polyimides.

The carbon fibers 86 a provide the material 86 with strength andstiffness. For example, the properties of the carbon fibers 86 a can beselected in accordance with desired properties of the material 86, andthus desired properties of the propulsor blades 74. In one example, thecarbon fibers 86 a are polyacrylonitrile or polyacrylonitrile-based. Thefibers are initially with polyacrylonitrile fibers and are thengraphitized. Alternatively, the fibers are initially thermoplasticfibers that are then graphitized. Thermoplastics can include, but arenot limited to, polyethylene, polyarylether, and poly ether ketones. Infurther examples, the carbon fibers 86 a have an average diameter of1-100 micrometers. Alternatively, the carbon fibers 86 a are nano-sizedand have a diameter of less than 1 micrometer. In other examples, thecarbon fibers 86 a are carbon-containing such that the fibers includecarbon as a primary constituent or element. In one example, the carbonfibers 86 a are carbide.

FIG. 6 illustrates a cross-sectional view of another example propulsorblade 174, which may include any of the aforementioned features. In thisdisclosure, like reference numerals designate like elements whereappropriate and reference numerals with the addition of one-hundred ormultiples thereof designate modified elements that are understood toincorporate the same features and benefits of the correspondingelements. In this example, the propulsor blade 174 includes a distinctcore 174 a that supports a skin 174 b of the carbon-fiber reinforcedpolymer matrix material 86. In this example, the core 174 a is a solidpiece, but it alternatively can be hollow to reduce weight.

The core 174 a can be formed of a metallic material, a fiber reinforcedpolymer matrix material, or combinations thereof. An example metallicmaterial includes a titanium-based alloy. The fiber reinforced polymermatrix material can include carbon fiber, as in any of the examples ofthe material 86. Alternatively, the fibers in the core 174 a arenon-carbon fibers. Example non-carbon fibers can include, but are notlimited to, glass fibers, metallic fibers, ceramic fibers, polymericfibers, and combinations thereof.

In further examples, the core 174 is formed of a fiber-reinforcedmaterial that is different in composition from the material 86 of theskin 174 b. The difference in composition can be in the kinds ofpolymers of the matrices, the kinds of fibers, the amounts of thepolymer matrices, the amounts of the fibers, or any combination of suchdifferences.

In further examples, the skin 174 b is the multi-layered structure ofthe material 86. For example, layers 88 are laid-up on or around thecore 174 a and then consolidated. Alternatively, the skin 174 b is acontinuous sleeve. The core 174 a is inserted into the sleeve and thenthe skin 174 b is consolidated. In one further example, the material 86of the sleeve has a three-dimensional fiber structure.

FIG. 7 illustrates another example propulsor blade 274 that is formed ofthe material 86. In this example, the propulsor blade 274 also includesa sheath 275 on a leading edge of the blade. For example, the sheath 275protects the propulsor blade 274 from foreign object impact. In oneexample, the sheath 275 is formed of a metallic material. The metallicmaterial can include, but is not limited to, a titanium-based alloy, acobalt-based alloy, or combinations thereof. In further examples, thesheath 275 is multi-layered and includes at least one layer of ametallic material. One or more additional layers can include a layer ofa metallic material of a different composition, a layer of apolymer-based material, or combinations thereof.

The sheath 275 is secured to the leading edge of the propulsor blade274. In this regard, the sheath 275 can be bonded using an adhesive,mechanically attached to the blade, or secured by a combination ofadhesive bonding and mechanical attachment.

In a further example, the propulsor blade 274 includes a first distinctregion 289 a (outside of dashed line region) of carbon-fiber reinforcedpolymer matrix material 86 and a second distinct region 289 b (insidedashed line region) of a non-carbon-fiber reinforced polymer matrixmaterial. The non-carbon fibers can include, but are not limited to,glass fibers, aramid fibers, boron fibers, carbide fibers, orcombinations thereof. The second distinct region 289 b ofnon-carbon-fiber reinforced polymer matrix material provides the abilityto locally tailor the performance of the propulsor blade 274 with regardto properties. For example, the vibrational properties are locallytailored through selection of the properties of the second distinctregion 289 b to control vibration or control response to an impactevent.

FIG. 8 illustrates selected portions of the fan section 22 of the engine20, including the case 43 and a portion of one of the propulsor blades74. The case 43 serves as a containment structure in the case of a bladerelease event. For example, the case 43 includes a fiber reinforcedpolymer matrix material 45. The material 45 includes fibers 45 a thatare disposed in a polymer matrix 45 b. The fibers 45 a can be carbonfibers or non-carbon fibers. Non-carbon fibers can include, but are notlimited to, glass fibers, aramid fibers, or combinations thereof. In oneexample, the material 45 includes a plurality of fiber layers 45 c thatare stacked and consolidated to form the material 45. For example, allof the layers 45 c have the same kind of fibers. In other examples,alternating layers 45 c, or an alternating pattern of layers 45 c, havedifferent kinds of fibers, one of which is carbon fibers.

In further examples, the carbon-fiber reinforced polymer matrix material86 of the propulsor blades 74 is different from the carbon-fiberreinforced polymer matrix material 45 of the case 43 with respect tocomposition. The difference in composition can be in the kinds ofpolymers of the matrices, the kinds of fibers, the amounts of thepolymer matrices, the amounts of the fibers, or any combination of suchdifferences. Further, the differences can be tailored for thermalconformance between the propulsor blades 74 and the case 43.

FIG. 9 illustrates another example case 143 that includes a layer of thematerial 45 adjacent a layer 147. The layer 147 can be a layer ofcarbon-fiber reinforced polymer matrix material, non-carbon-fiberreinforced polymer matrix material, or metallic material, such as in ahoneycomb or acoustic structure.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. The scope of legal protection given tothis disclosure can only be determined by studying the following claims.

What is claimed is:
 1. A gas turbine engine comprising: a gear assembly; a bypass flow passage; a fan located upstream of the bypass flow passage; a first shaft and a second shaft; a first turbine coupled through the gear assembly to the fan; a first compressor coupled with the first shaft; and a second turbine coupled with the second shaft; wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 16.0.
 2. The gas turbine engine as recited in claim 1, wherein the number of fan blades is 20 and N/R is 15.4.
 3. The gas turbine engine as recited in claim 1, wherein the number of fan blades is 18 or
 20. 4. The gas turbine engine as recited in claim 3, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio is from 1.2 to 1.4.
 5. The gas turbine engine as recited in claim 3, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio is from 1.3 to 1.4.
 6. The gas turbine engine as recited in claim 1, further comprising a case surrounding the fan, and the fan is a single fan stage.
 7. The gas turbine engine as recited in claim 6, wherein each of the fan blades is fixed in position between the hub and the tip.
 8. The gas turbine engine as recited in claim 7, wherein the fan blades are formed of a fiber-reinforced polymer matrix material.
 9. The gas turbine engine as recited in claim 6, wherein the first compressor is a 3-stage compressor.
 10. The gas turbine engine as recited in claim 7, wherein the first turbine is a 5-stage turbine.
 11. The gas turbine engine as recited in claim 10, further comprising a second turbine coupled with the second shaft, and the second turbine is a 2-stage turbine.
 12. A gas turbine engine comprising: a gear assembly; a bypass flow passage; a fan located upstream of the bypass flow passage; a first shaft and a second shaft; a first turbine coupled through the gear assembly to the fan; a first compressor coupled with the first shaft; and a second turbine coupled with the second shaft; wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 13.8 to 16.4.
 13. The gas turbine engine as recited in claim 12, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, the design pressure ratio is from 1.2 to 1.4.
 14. The gas turbine engine as recited in claim 12, wherein each of the fan blades is fixed in position between the hub and the tip, further comprising a case surrounding the fan, the fan is a single fan stage, and the fan blades are formed of a fiber-reinforced polymer matrix material.
 15. The gas turbine engine as recited in claim 14, wherein the first compressor is a 3-stage compressor and the first turbine is a 5-stage turbine, further comprising a second turbine coupled with the second shaft, and the second turbine is a 2-stage turbine.
 16. A gas turbine engine comprising: a gear assembly; a bypass flow passage; a fan located upstream of the bypass flow passage; a first shaft and a second shaft; a first turbine coupled through the gear assembly to the fan; a first compressor coupled with the first shaft; and a second turbine coupled with the second shaft; wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is 18 or 20 and a solidity value (R) at tips of the fan blades, where for 20 fan blades the solidity value (R) is 1.3, 1.1, or 1.0 such that a ratio of N/R is, respectively, 15.4, 18.2, or 20, or where for 18 fan blades the solidity value (R) is 1.3, 1.2, or 1.1 such that the ratio of N/R is, respectively, 13.8, 15, or 16.4.
 17. The gas turbine engine as recited in claim 16, wherein the number of fan blades is
 20. 18. The gas turbine engine as recited in claim 16, wherein the number of fan blades is
 18. 19. A gas turbine engine comprising: a bypass flow passage; a fan located upstream of the bypass flow passage; a first turbine; a first compressor; and means for coupling an extraction of energy from the first turbine to the first compressor and the fan, wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to
 16. 20. The gas turbine engine as recited in claim 19, wherein the means for coupling an extraction of energy comprises a first shaft, wherein the first turbine is coupled with the first shaft, and the first shaft is coupled with the first compressor and the fan is coupled to a gear assembly which is coupled to the first shaft. 